Method for manufacturing composite parts

ABSTRACT

A method for manufacturing composite parts includes positioning a plurality of composite layers on an application surface formed by a plurality of tooling members and filler members on a mandrel. A part formation aid is disposed on the composite layers above each tooling member. The filler members are then removed, and the composite layers slit over the recesses. The mandrel, the tooling members, the composite layers, and the part formation aids are placed in a vacuum bag. As the bag is evacuated, each part formation aid forces the side edge of the composite layers formed by the slit around the respective tooling member, thereby forming a composite part with a desired shape, thickness, and density. In an alternate embodiment, a flanged composite panel is formed around a rectangular mandrel after positioning composite layers on an outer surface of the mandrel and an outer surface of a filler member disposed on one end of the mandrel. After removing the filler member, the flanged panel is formed and partially cured by vacuum bagging. A plurality of tooling members and filler members are then placed on the flanged panel to form an application surface. Composite layers are positioned on the application surface, and then the filler members are removed. Vacuum bagging is used to produce a plurality of stiffeners around the tooling members and directly adjacent to the flanged panel. The stiffeners and flanged panel are co-cured to produce a single and integral part.

This application is a continuation of application Ser. No. 09/591,352,filed Jun. 9, 2000, now U.S. Pat. No. 6,406,580, entitled “Method forManufacturing Composite Parts”.

BACKGROUND OF THE INVENTION

1. Field of the Invention

This invention relates in general to a method of manufacturingcomposites and in particular to a method for manufacturing compositepanels and stiffeners by first preparing a circumferential array ofmaterial and then forming and curing the material to the desired shape.

2. Description of Related Art

Composite structures and parts are necessary parts for industriesrequiring high strength, lightweight materials. A good example of thisis in the aerospace industry, where aircraft and other airborne vehiclesrequire high strength components that weigh as little as possible.

Many approaches have been previously developed for forming multiplelayers of composite material into a desired shape or shapes. The mostcommon, particularly in the aircraft industry, involves placingindividual layers of material onto a form having a desired shape, andthen curing the layers. Curing the material through application of heatand pressure fully compacts or debulks the composite material. The curedcomposite material then has the desired shape and strength. Formingparts in this way does not involve significant reshaping of thecomposite material during curing and may be very time consuming.

Disadvantages inherent in the aforementioned process include the verytedious and time consuming operation of laying individual layers ofcomposite material directly onto a tool to obtain a final non-flatdesired shape. The very labor intensive process of placing the layers ofmaterial onto a form may require many highly-skilled man-hours for eachpart, and is, therefore, very expensive. Additionally, theaforementioned process may require stopping after placement of every fewlayers of material and providing some form of mechanical compaction tothe material. This may be necessary to achieve final full compaction ofthe layers. Failure to achieve full compaction of the material layersprior to curing may result in wrinkles and other anomalies in the finalstructure, since as individual layers compact, the local path-lengths ofthe fibers in the layers change. Wrinkles and other anomalies in thecured structure are aesthetically and structurally undesirable.

Previously developed methods for forming composite parts also fail toassure uniformity between parts. In many prior art methods, each part isseparately made. Each part is formed by the process of placingindividual layers onto a form and then curing the layers while on theform. The cured part is removed from the form, allowing the next part tobe made by the same process. By this method a number of parts can beformed. Unfortunately, variations in compaction, in resin bleeding fromthe part, and in fiber “washing” or dislocations from resin bleeding,tend to occur because compaction is occurring three dimensionally, andbecause of the low viscosity of the resin. These factors may yield partsthat lack uniformity. Previously developed methods for buildingcomposite parts are, therefore, not compatible with low-cost,high-volume manufacturing methodologies.

Composite parts fabricated by previously developed methods often requiremachining after curing, e.g. routing, grinding, etc., in order to meetfinal dimension requirements. This machining adds additional time andexpense to the process of fabricating the part and can result in damageto the part by delamination of the cured layers.

Yet another disadvantage of the previously developed methods forfabricating composite parts is their incompatibility with in-processcontrol (IPC), statistical process control (SPC), and total quality (TQ)methodologies. IPC, SPC, and TQ require repeatable, measurable resultsto obtain full effectiveness. The custom approach of the prior art tofabricating composite parts is not amenable to obtaining the benefits ofIPC, SPC, and TQ, i.e., high quality, high yield, and low cost.

Previously developed methods for forming composite parts often do notprovide acceptable results when forming complex parts from two or moresub-parts or pre-forms by “co-curing”. In co-curing, two or moresub-parts are made into a single part by placing the sub-parts in thedesired orientation and curing the combination. Since the prior artrequires the individual layers of a sub-part to be laid-up in theirfinal shape on a form joining two or more individual sub-parts to make apart, e.g., two channels and two plates to form an I-beam, is verydifficult. If a foreign material, e.g., backing paper or tape, isaccidentally trapped between the layers during layup of a part, there islittle likelihood that it will be detected. As a result, high laborcosts may be invested in a complex, co-cured part that must be scrappeddue to the inclusion.

Prior methods for fabricating composite parts often require that theindividual composite layers be stored in a freezer prior to layup. Thisadds additional handling and equipment costs in fabricating a compositepart.

Therefore a need has arisen for an improved method and system forfabricating parts from composite materials.

A need further exists for an improved method and system for reducing thetime necessary for fabricating composite parts.

A further need exists for a low-cost method and system for fabricatingparts from composite materials.

Yet another need exists for a method and system for fabricating multipleuniform parts from composite materials that do not require significantamounts of machining after curing.

Another need exists for a method for fabricating composite partscompatible with IPC, SPC, and TQ.

Yet another need exists for a method and system for fabricatingcomposite parts that eliminate the need for special handling and storageof composite layers.

BRIEF SUMMARY OF THE INVENTION

The present invention provides a method for manufacturing a compositepart having a specified shape, thickness, and density. The compositepart is formed in an array of similar parts located around a mandrelhaving an outer surface and a longitudinal axis. Several tooling membersare disposed around the outer surface of the mandrel, the toolingmembers serving as the molds around which the composite parts will beformed. A plurality of filler members are disposed on the outer surfaceof the mandrel between the tooling members. An outer surface of eachtooling member and an outer surface of each filler member combine toform a generally smooth and rounded application surface that surroundsthe mandrel.

A plurality of composite layers is positioned on the application surfaceusing an automated positioning technique such as fiber placement. A partformation aid is placed on the composite layers above each toolingmember, and a cut is made in the composite layers parallel to thelongitudinal axis of the mandrel and between each tooling member. Thefiller members are then removed from between the tooling members.

Finally, the mandrel and composite layers are placed in a vacuum bag,which is then placed in an autoclave. The vacuum bag is an elastomericmaterial in the form of a tube that is slid over the mandrel. The vacuumbag tube can be pulled out of the way during fiber placement and slidover the mandrel at the time of forming. As the bag is evacuated, eachpart formation aid deforms toward the mandrel, thereby forming thecomposite layers around the tooling members. The composite partresulting from the evacuation of the bag is then cured for a specifiedamount of time to insure that the composite part will maintain itsspecified shape, thickness, and density.

Alternatively, the method according to the present invention is used tofabricate composite parts having more than one component. Oneapplication of this method is to produce a flanged panel having aplurality of hat-shaped stiffeners integrally disposed on the panel.

The flanged panel is produced using a rectangular mandrel having anouter surface. A filler member having an outer surface is placed on oneend of the mandrel, the outer surface of the filler member and the outersurface of the mandrel forming a generally smooth and roundedapplication surface.

A plurality of composite layers is positioned on the application surfaceusing an automated positioning method such as fiber placement. Afterplacement of the layers, the filler member is removed, and the compositelayers are subjected to the vacuum bagging technique previouslymentioned. This process forms a flanged composite panel. The panel isthen partially cured.

After partial curing, a plurality of tooling members are disposed on anouter surface of the recently formed panel. Filler members are placed onthe outer surface of the panel between the tooling members to form agenerally smooth second application surface. Composite layers areapplied to the second application surface using fiber placement. Thefiller members are then removed and the composite layers are cut betweenthe tooling members.

The flanged panel, the tooling members, and the newly applied compositelayers are placed in a vacuum bag. As the bag is evacuated, thecomposite layers form into hat-shaped stiffeners around the toolingmembers. The flanged panels and the stiffeners are finally co-cured orco-bonded to form an integral composite part.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a partially cut-away perspective view showing a plurality ofcomposite layers positioned around a mandrel, tooling members, andfiller members according to the present invention.

FIG. 2 is a perspective view showing the mandrel, tooling members, andcomposite layers of FIG. 1 with a part formation aid located above eachtooling member.

FIG. 3 is a partially cut-away perspective view of the mandrel andtooling members of FIG. 1 shown installed in a vacuum bag, the vacuumbag being held to the inside of a hard outer tube.

FIG. 4 is a sectional view of one of the tooling members, shown duringevacuation.

FIG. 5 is a perspective of view of a composite part formed using themethod of the present invention with the mandrel and tooling members ofFIG. 1.

FIG. 6 is a perspective view of a mandrel having a plurality ofhat-shaped tooling members for use with the method of the presentinvention.

FIG. 7 is a front view of a mandrel having a plurality of Z-shapedtooling members for use with the method of the present invention.

FIG. 8 is a perspective view of a composite part produced by using analternate embodiment of the method according to the present invention.

FIG. 9 is a perspective view of a plurality of composite fiberspositioned on a rectangular mandrel and a filling member according tothe alternate embodiment of the present invention.

FIG. 10 is a perspective view of a plurality of composite fiberspositioned on a plurality of tooling members installed on a compositepart produced from the composite fibers of FIG. 9.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

Referring to FIGS. 1, 2, and 3 in the drawings, a method ofmanufacturing composite parts according to the present invention isillustrated. A mandrel 11 having a longitudinal axis 13 and an outersurface 15 provides a support base for producing the composite partsaccording to the present invention. In this embodiment, mandrel 11 issquare in cross section, having four flat sides on its outer surface 15.A plurality of tooling members 17 are disposed on outer surface 15 ofmandrel 11. Tooling members 17 generally have a longitudinal axis thatis positioned parallel to longitudinal axis 13 when tooling members 17are attached to mandrel 11. Although tooling members 17 are generallymade of rubber and are mechanically fastened to mandrel 11, it isconceivable that tooling members 17 could be an integral part of mandrel11. In cross-section, each tooling member is a trapezoid in thisembodiment, having sidewalls 18 that converge toward each other and anouter surface 25 that is a portion of a cylinder.

After installation of tooling members 17, filler members 21 are placedbetween tooling members 17 and against outer surface 15 of mandrel 11.Depending on the spacing between tooling members 17 and thecross-sectional shape of mandrel 11, more than one filler member 21 maybe required in the space between tooling members 17. In FIG. 1, twofiller members 21 are placed in each space between tooling members 17.Each filler member 21 has a sidewall 22 that diverges from outer surface15 and abuts one of the sidewalls 18. Each filler member 21 has an outersurface 27 that is a portion of a cylinder.

Filler members 21 are used in conjunction with tooling members 17 toprovide a smooth application surface 23. Typically, filler members 21are made of tooling foam, which is easily moldable or shapable intoapplication surface 23. Tooling foam was used because of cost andbecause this was a one time part. Materials for the filler members 21are chosen for production with durability as a requirement as well ascost or because of ease of manufacture. Application surface 23 is formedby outer surface 25 of each tooling member and outer surface 27 of eachfiller member 21, which are flush. Usually, application surface 23 willbe smooth and rounded, and preferably, application surface 23 will havea cylindrical shape as shown in FIG. 1.

After preparing application surface 23, a plurality of composite layers37 (not shown in FIG. 3) are positioned on the application surface 23.Several different methods could be used to position composite layers 37on application surface 23. The most basic method would be to hand laythe composite material. However, a major advantage of the method of thepresent invention is its adaptability towards automation. Although anyautomated technique for applying composites could be used, the mostcommon automated techniques include filament winding, tape winding, andfiber placement.

Filament winding is a process in which a continuous filament, usuallyconstructed of a reinforced fiber impregnated by a matrix material, iswound under tension around a rotating core. Referring to FIG. 1, mandrel11 is rotated as the filament is applied under tension to applicationsurface 23. Several rotations of mandrel 11 would be necessary tocompletely apply the filament, thus constructing the plurality ofcomposite layers 37. The matrix material is either preimpregnated in thereinforced fiber or is applied to the fiber as the filament is beingpositioned on application surface 23. The matrix material serves toadhere the filament to previous layers as it is being applied, and alsoserves as a curing agent after formation of a final composite part.

A similar automated application technique is tape winding. A tow is abundle of more than a thousand filaments. In a tape winding process, aunidirectional preimpregnated tape is used that consists of several towsthat have been formed and spread with resin. The tape is cut to aparticular width, common widths being 0.5 inch and 1.0 inch. The tape ispositioned on application surface 23 as mandrel 11 is rotated aboutlongitudinal axis 13. The tape is applied under tension, and no contactis made between the head of the winding machine and application surface23.

The preferred application technique for the present invention is fiberplacement. Unlike filament winding and tape winding, fiber placementdoes not use rotation of mandrel 11 to pull the composite material ontoapplication surface 23. Instead, a fiber placement machine having a headmakes contact with application surface 23 and is used to apply either atow or a small width tape to the application surface 23. As material isplaced on application surface 23, more tows or tapes are pushed to thehead of the fiber placement machine. The head of the machine applies aforce to the tow or tape as it is applied. The primary advantage ofusing fiber placement is that the directional placement of the materialis not limited by the rotation of mandrel 11. Instead, material can beplaced in any direction. Additionally, scrap is reduced when comparedwith filament winding and tape winding methods. Finally, fiber placementallows material to be easily added or removed before the material isformed into a final composite part.

After applying composite layers 37 to application surface 23, a partformation aid 39 optionally may be installed on an outer surface ofcomposite layers 37 above each tooling member 17. Part formation aids39, also referred to as “bat-wings,” aid in the formation of a finalcomposite part and are shaped based on the final part to be formed. Inthis embodiment, part formation aid 39 is generally U-shaped, having abase 40 and two outward extending sidewalls 42. The part formation aid39 shown in FIGS. 2 and 3 is used to assist in the formation of ahat-shaped stiffener. Part formation aids 39 are sufficiently flexibleto allow the sidewalls 42 to flex inward into an inverted shape to thatshown in FIG. 2.

After installation of part formation aids 39, composite layers 37 arecut parallel to longitudinal axis 13 and between each tooling member 17along slit 44 (FIG. 2). The cutting of the composite material 37 allowsindividual parts to be formed around each tooling member 17. Slit 44results in a plurality of separate composite layer segments, each havingside portions extending in opposite directions from each tooling member17. For production it is recommended that the cutting operation beautomated. The automation can be accomplished by attaching a cuttingblade to a multiple axis end effector of the fiber placement machine orwinding machine. Alternatively, fiber placement machines have an abilityto stop and start material placement on the fly. Therefore, the machinecould be programmed to leave a gap in the material between the segments.

Filler members 21 (not shown in FIG. 2) are then removed from mandrel11, thus forming recesses between the tooling members 17. Mandrel 11,which still has tooling members 17, composite layers 37, and partformation aids 39 attached, is placed in a vacuum bag 45 as illustratedin FIG. 3. The vacuum bag tube 45 is placed over the mandrel with theaid of a hard outer tube. The hard outer tube is made in an appropriateshape to allow for easy fit over the mandrel 11. Vacuum bag 45 is usedin conjunction with two sealing end plates 47, a vacuum fitting 49, anda vacuum line 51. After placing vacuum bag 45 over mandrel 11, the bag45 is sealed to each end plate 47 using a typical band clamp 53. Vacuumline 51 is then connected at one end to vacuum fitting 49 and at itsother end to a vacuum source (not shown).

The formation of a composite part using the present invention isaccomplished by creating a vacuum within vacuum bag 45. After vacuum bag45 has been sealed, the air from the bag can be evacuated through vacuumline 51 leading from a central passage in mandrel 11. The evacuation ofair causes bag 45 to apply a formation force to each part formation aid39 that is directed toward the center of mandrel 11. As the formationforce is applied to each part formation aid 39, the sidewalls 42 of partformation aid 39 deform downward (toward the center of mandrel 11),thereby causing the side portions of the segments of composite layers 37to be forced down and around tooling member 17. The folding sidewalls 42cause side portions of composite layers 37 alongside each slit 44 to bepushed into abutment with one of the sidewalls 18, as shown in FIG. 4.Tooling members 17 define the general shape that a composite part 61(FIG. 5) produced by the current process will become. After initialformation of the composite part, the part is held in place by vacuum bag45 during a curing period.

Although not illustrated, the forming and curing processes describedabove are usually carried out inside an oven or autoclave. An autoclaveallows the ambient temperature and pressure surrounding vacuum bag 45 tobe raised. The exact temperature and pressure requirements for formingand curing a part depend on the resin that is used with the compositelayers 37. Typically, curing takes place at elevated temperatures andpressures. Some resins actually allow the forming and curing processesto take place at room temperature and atmospheric pressure, which wouldnegate the need for an autoclave.

After forming and curing, the composite part 61 (FIG. 5) is in its finalshape. In FIGS. 1, 2, and 3, the shape of tooling members 17 is designedto produce hat-shape stiffener 61 as shown in FIG. 5. Stiffener 61 iscommonly used with flat panels to provide added rigidity and strength tothe panels. The most common application of composite stiffener 61 is inaerospace applications where strong and lightweight materials areneeded.

Mandrel 11 is configured to produce four hat-shape stiffeners 61 perproduction run. Referring to FIG. 6, a mandrel 71 having a cylindricalouter surface 73 with a plurality of tooling members 75 is illustrated.Because of the large number of tooling members 75 disposed on mandrel71, the production of stiffeners 61 would be greatly increased over theconfiguration provided by mandrel 11. The process described above forforming stiffeners 61 using mandrel 11 is the same process that would beused to produce stiffeners 61 using mandrel 71.

Referring to FIG. 7, a mandrel 81 having integrally formed toolingmembers 83 is illustrated. The shape of tooling members 83 is such thata stiffener 85 having a Z-shaped cross section is formed using theprocess described above. Foam blocks (not shown) would be initiallylocated between each tooling member 83, resulting in a cylindricalexterior. Fibers would placed over the cylindrical exterior. The foamblocks would be removed. Formation aids would bend the fibers into theshape as shown in FIG. 7. Although tooling member 83 are shown as anintegral portion of mandrel 81, tooling members 83 could be separateparts that are mechanically fastened to the mandrel 81. Again, amaterial other than foam could be used in place of the foam blocks.

Referring to FIGS. 8, 9 and 10, an alternate method of manufacturingcomposite parts according to the present invention is illustrated. Acomposite part 111 (FIG. 8) having a first portion, or flanged panel 113and a plurality of second portions, or stiffeners 115 is produced, firstportion 113 being co-cured to second portions 115. Composite part 111has a flanged composite panel (first portion 113) with severalhat-shaped stiffeners (second portions 115) attached to the panel.

Referring more specifically to FIG. 9, a mandrel 117 having an outersurface 119 is provided to aid in the formation of first portion 113.Although the actual shape of first portion 113 and thus the shape of themandrel could vary depending on the application, the flanged compositepanel 113 of part 111 is produced using a rectangular mandrel 117.

Before forming flanged panel 113, a filler member 121 is attached toouter surface 119 of mandrel 117 in any place needed to form a roundedor smooth corner for the formation of first portion 113. An outersurface of filler member 121 combines with outer surface 119 to form anapplication surface 123.

A plurality of composite layers 131 are positioned on applicationsurface 123 using fiber placement, tape winding, filament winding, orhand-laying methods. A preferred method of applying composite layers 131is fiber placement. Fiber placement of the composite material 131 wouldprevent having to wrap composite layers 131 around the entirecircumference of mandrel 117. Instead the composite layers 131 can beplaced wherever material is needed on application surface 123.

After applying composite material 131, the material 131 is cut alongslit 133 to trim and remove any excess material. Filler member 121 isthen removed from the mandrel 117. Mandrel 117 and composite material131 is placed in a vacuum bag (not shown) as described previously. Thevacuum bag is sealed and evacuated of air, causing the bag to apply aforce to the composite layers 131, thereby forming the flanged compositepanel 113 of composite part 111.

Depending on the curing requirements of the composite material 131 used,the evacuation process may or may not take place inside an autoclave.Regardless of whether an autoclave is used, flanged panel 113 is only tobe partially cured. This allows stiffeners 115 to be co-cured to flangedpanel 113.

Referring more specifically to FIG. 10, a plurality of tooling members141, each having a center channel 142, is disposed on a surface offlanged panel 113 opposite mandrel 117. The shape of tooling members 141determine the final shape of stiffeners 115. In FIG. 8, tooling members141 are trapezoidal in cross-section to produce a hat-shaped stiffenersimilar to stiffener 61. Filler members 143 are placed on flanged panel113 between tooling members 141 to form a smooth flat applicationsurface 145.

A plurality of composite layers 151 are positioned on applicationsurface 145. The preferred method of applying composite material 151 isfiber placement, although any method of positioning composites could beused. Part formation aids (not shown but similar to part formation aids39) are placed on composite layers 151 above each tooling member 141.Composite layers 151 are cut between each tooling member 141 on a lineparallel to the lengthwise axis of each tooling member 141. Fillermembers 143 are removed from between tooling members 141.

Stiffeners 115 are formed using the evacuation technique previouslydescribed. Mandrel 117, first portion 113, tooling members 141,composite layers 151, and the part formation aids are placed inside avacuum bag (not shown). The vacuum bag is sealed and evacuated, therebycausing the bag to exert a force on the part formation aids. The partformation aids deform toward flanged panel 113, which causes compositelayers 151 to conform to the shape of tooling members 141.

After forming stiffeners 115, both the flanged panel 113 and thestiffeners 115 are co-cured to obtain the composite part 111. Dependingon the requirements of the resin used in composite layers 131, 151, theforming and curing processes will likely be performed inside anautoclave at elevated temperatures and pressures.

After forming composite part 111, tooling members 141 are removed bycreating a vacuum within center channel 142. Since tooling members 141are generally made from rubber, the tooling members 141 will compressinwardly and are easily removed from within stiffeners 115.

The scope of the alternate method described is not limited to forming apart having only a first portion and a second portion. Instead, a parthaving a plurality of portions may be formed by similar steps, eachportion being co-cured to another in one of the final steps. Finally, itis conceivable that a similar method be used to form and attach acomposite part to a preexisting composite part formed by a processoutside the scope of the present invention.

One advantage of the present invention is that it provides an automatedmethod for manufacturing composite parts using a circular array oftooling members. By providing a smooth and partially rounded applicationsurface, composite material can be quickly and efficiently positioned byan automated positioning technique (i.e. fiber placement, tape winding,filament winding). Additionally, the circular configuration of toolingmembers allows rapid, large-scale production of complex composite shapesthat normally require extensive manufacturing time.

Another advantage of the present invention is that it provides a methodof automating the manufacture of two or more composite components thatwill form a final composite part. By preparing smooth applicationsurfaces, automated composite positioning techniques can be used tocreate any number of different composite components. The variouscomponents can then be joined to form the final part during a co-curingor co-bonding process.

It should be apparent from the foregoing that an invention havingsignificant advantages has been provided. While the invention is shownin only a few of its forms, it is not just limited but is susceptible tovarious changes and modifications without departing from the spiritthereof.

I claim:
 1. A method of manufacturing a composite part, comprising: (a)positioning a plurality of spaced-apart tooling members on a supportingsurface; (b) placing filler members between the tooling members, thefiller members and the tooling members having flush outer surfaces suchthat a generally smooth application surface is formed by the outersurfaces of the tooling members and the filler members; (c) providingcomposite layer segments on the application surface; (d) removing thefiller members, creating recesses between each of the tooling members,each of the composite layer segments having a side edge portionextending over one of the recesses; (e) forming the composite layersegments into the composite part's final shape by folding the sideportions into the recesses into abutment with each of the toolingmembers; and (f) holding the composite layer segments in the part'sfinal shape until sufficiently cured to maintain the part's final shape.2. The method according to claim 1 wherein the generally smoothapplication surface is a cylindrical surface.
 3. The method according toclaim 1 wherein step (f) comprises applying heat.
 4. The methodaccording to claim 1 wherein steps (e) and (f) comprise: inserting thesupporting surface, tooling members, and composite layer into a vacuumbag, then placing the vacuum bag into an oven and drawing a drawing avacuum within the vacuum bag.
 5. The method according to claim 1 whereinthe composite layer is applied to the application surface using a fiberplacement apparatus.
 6. The method according to claim 1 wherein thecomposite layer is applied to the application surface by windingfilaments of the composite layer around the application surface.
 7. Themethod according to claim 1 wherein the supporting surface comprises amandrel.
 8. The method according to claim 1 wherein the supportingsurface comprises a flat composite panel, and step (f) further comprisesbonding the side edge portion of the composite layer segments to thecomposite panel.
 9. The method according to claim 1 further comprising:placing a flexible part formation aid on the composite layer above theouter surface of each tooling member before step (e), the part formationaid deflecting during step (e) to physically force the side edgeportions of each composite layer segment into the recess.
 10. The methodaccording to claim 1 further comprising rotating the mandrel about alongitudinal axis during step (e).
 11. The method according to claim 1wherein the composite layer is made of preimpregnated fiber.
 12. Amethod of manufacturing composite parts, comprising: providing acomposite first portion having an outer surface; attaching a pluralityof tooling members to the outer surface of the first portion; placingfiller members adjacent to the outer surface of the first portion andbetween the tooling members, the filler members being flush with thetooling members such that a generally smooth application surface isformed; providing composite layer segments on the application surface;removing the filler members, defining recesses, each of the compositelayer segments having a side edge portion extending over one of therecesses; forming composite second portions into their final shape bydeflecting the side edge portions into the recesses; and holding thesecond portions in the final shape until the second portions aresufficiently bonded to the first portion such that both first and secondportions are joined and both first and second portions maintain theirfinal shape.
 13. The method according to claim 12 wherein the step ofproviding a composite first portion further comprises: providing amandrel having at least one surface in the shape of the first portion tobe formed; attaching a filler member to the mandrel to establish agenerally smooth application surface for the first portion that isformed by an outer surface of the filler member and the at least onesurface of the mandrel; applying a composite layer to the applicationsurface; removing the filler member from the mandrel to create a recessand cutting the composite layer along the recess such that the compositelayer has at least one edge portion extending over the recess;deflecting the edge portion into the recess, thereby forming the firstportion into its final shape; and holding the first portion in the finalshape until the first portion is partially cured.
 14. The methodaccording to claim 12 wherein the first portion formed comprises a panelhaving a flange.
 15. The method according to claim 12 wherein the secondportions formed comprise elongated stiffener members.